A turbine blade of a turbine, which is mounted on a jet engine or the like, is provided with various countermeasures against heat due to its exposure to a high-temperature gas such as a combustion gas generated at a combustor.
These countermeasures include cooling methods to suppress the overheating of the turbine blade by forming a hollow region in the turbine blade and supplying a coolant gas to the hollow region. One of those cooling methods is an impingement cooling method. The impingement cooling method has been known as a cooling method by which high cooling performance is obtained. However, since parts called inserts need to be inserted into a blade, the shape of the blade is limited when the impingement cooling method is used. In the present aerodynamic design of a blade, the blade generally has a complicated three-dimensional shape in order to improve blade element performance. For this reason, the impingement cooling method restricts the aerodynamic design of a blade since the shape of a blade is limited so that inserts can be inserted into the blade.
A technique disclosed in WO 2007/094212 A1, (page 19, FIG. 10), is proposed as a technique that delivers the same cooling performance as the cooling performance of the impingement cooling method while making up for the shortcomings of the impingement cooling method.
Specifically, a structure, which improves cooling efficiency by causing a coolant gas to repeatedly meander between a suction wall surface and a pressure wall surface while guiding the coolant gas toward a trailing edge side from a leading edge side of a turbine blade, is disclosed in WO 2007/094212 A1, (page 19, FIG. 10).